Lost core structural frame

ABSTRACT

A lost core mold component comprises a first leg and a second leg with a plurality of crossover members connecting the first and second legs. The plurality of crossover members includes outermost crossover members spaced from each other. Adjacent ends of each of the first and second legs, and second crossover members are spaced closer to each other than are the outermost crossover members. Central crossover members extend between the first and second leg and between the second crossover members. The outermost crossover members extend for a first cross-sectional area. The second crossover members extend for a second cross-sectional area and the central crossover members extend for a third cross-sectional area. The first cross-sectional area is greater than the second cross-sectional area. The second cross-sectional area is greater than the third cross-sectional area. A gas turbine engine and component are also disclosed.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.15/039,918, filed May 27, 2016, now U.S. Pat. No. 10,370,980 grantedAug. 6, 2019, which is a National Phase of International Application No.PCT/US2014/066495, filed Nov. 20, 2014 which claims priority to UnitedStates Provisional Patent Application No. 61/919,846, filed Dec. 23,2013.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under Contract No.N00019-02-C-3003, awarded by the United States Navy. The Government hascertain rights in this invention.

BACKGROUND OF THE INVENTION

This application relates to a lost core for forming cooling channels ina gas turbine engine component airfoil.

Gas turbine engines are known and, typically, include a fan deliveringair into a bypass duct as propulsion air and into a core engine, whereit reaches a compressor. The air is compressed and delivered into acombustion section. The air is mixed with fuel and ignited and productsof this combustion pass downstream over turbine rotors driving them torotate.

The products of combustion are extremely hot and, thus, there arechallenges to airfoil durability in the turbine section. As an example,the turbine rotors carrying rotating blades that have airfoils. Inaddition, there are static vanes having airfoils intermediate stages ofthe rotating blades. It is known to provide cooling air to internalchannels within these airfoils.

An airfoil typically extends from a leading edge to a trailing edge.Cooling channels are provided at the trailing and leading edges.

Internal cooling channels may be formed by a lost core mold techniques.In such technique, a lost core is made which essentially follows thecontours of the desired cooling channels. That lost core is placedwithin a mold and molten metal is molded around the lost core. The lostcore is then leached out leaving a hollow cavity within the solidifiedmetal airfoil.

In one application of lost cores, there is a first leg forming a coolingchannel adjacent to at least one of the leading and trailing edges.Crossover holes connect this channel to a channel spaced from theleading or trailing edge. These crossover holes are formed by crossovermembers on the core. To prevent breakage to the core during the moldingprocess and leading up to the molding process, the extreme end crossovermembers are formed to be of greater cross-sectional area. Such crossovermembers are called frames.

However, the frames have not always provided sufficient strength andprevention of damage.

SUMMARY OF THE INVENTION

In a featured embodiment, a lost core mold component comprises a firstleg and a second leg with a plurality of crossover members connectingthe first and second legs. The plurality of crossover members includesoutermost crossover members spaced from each other. Adjacent ends ofeach of the first and second legs, and second crossover members arespaced closer to each other than are the outermost crossover members.Central crossover members extend between the first and second leg andbetween the second crossover members. The outermost crossover membersextend for a first cross-sectional area. The second crossover membersextend for a second cross-sectional area and the central crossovermembers extend for a third cross-sectional area. The firstcross-sectional area is greater than the second cross-sectional area.The second cross-sectional area is greater than the thirdcross-sectional area.

In another embodiment according to the previous embodiment, a ratio ofthe first cross-sectional area to the second cross-sectional area to thethird cross-sectional area is 3:2:1.

In another embodiment according to any of the previous embodiments, thelost core member is further provided with pins to form film coolingholes.

In another embodiment according to any of the previous embodiments, thelost core member is utilized to form two cooling channels adjacent atleast one of a leading and trailing edge in an airfoil which is to bemolded around the lost core member.

In another embodiment according to any of the previous embodiments, thelost core member is utilized to form two cooling channels adjacent theleading edge.

In another embodiment according to any of the previous embodiments,there are intermediate hollows between the outermost crossover membersand the second crossover members, and between the second members and thecentral crossover members, and between individual ones of the centralcrossover members.

In another embodiment according to any of the previous embodiments, thehollows extend for a cross-sectional area that is less than the firstand the second cross-sectional area.

In another embodiment according to any of the previous embodiments, thelost core member is further provided with pins to form film coolingholes.

In another embodiment according to any of the previous embodiments, thelost core member is utilized to form two cooling channels adjacent atleast one of a leading and trailing edge in an airfoil which is to bemolded around the lost core member.

In another featured embodiment, a gas turbine engine component having anairfoil comprises an airfoil extending between a leading edge and atrailing edge. There is a first and second cooling channel with thefirst cooling channel being spaced closest to one of the leading andtrailing edges and the second channel being spaced from the firstchannel relative to the one of the leading and trailing edges. Crossoverholes connect the first and second cooling channels. The plurality ofcrossover holes include outermost crossover holes spaced from eachother. Adjacent ends of each of the first and second cooling channels,and second crossover holes are spaced closer to each other than are theoutermost crossover holes. Central crossover holes extend between thefirst and second cooling channels and between the second crossoverholes. The outermost crossover holes extend for a first cross-sectionalarea. The second crossover holes extend for a second cross-sectionalarea and the central crossover holes extend for a third cross-sectionalarea. The first cross-sectional area is greater than the secondcross-sectional area and the second cross-sectional area is greater thanthe third cross-sectional area.

In another embodiment according to the previous embodiment, a ratio ofthe first cross-sectional area to the second cross-sectional area to thethird cross-sectional area is 3:2:1.

In another embodiment according to any of the previous embodiments, filmcooling holes extend from the first cooling channel through a skin ofthe component.

In another embodiment according to any of the previous embodiments, atleast one of the leading and trailing edge is the leading edge.

In another embodiment according to any of the previous embodiments,there are intermediate connectors between the outermost crossover holesand the crossover holes, and between the second crossover holes and thecentral crossover holes, and also between individual ones of the centralcrossover holes.

In another embodiment according to any of the previous embodiments, theconnectors extend for a cross-sectional area that is less than the firstand the second cross-sectional area.

In another embodiment according to any of the previous embodiments, agas turbine engine comprises a compressor section and a turbine section.The turbine section includes rotors carrying rotating blades and staticairfoils with at least one of the rotating blades. The static airfoilsinclude an airfoil extending between a leading edge and a trailing edge.There is a first and second cooling channel with the first coolingchannel being spaced closest to one of the leading and trailing edges.The second channel is spaced from the first channel relative to the oneof the leading and trailing edges. Crossover holes are formed betweenthe first and second cooling channels. A first leg and a second leg witha plurality of crossover members connect the first and second legs. Theplurality of crossover members include outermost crossover membersspaced from each other. Adjacent ends of each of the first and secondlegs, and second crossover members are spaced closer to each other thanare the outermost crossover members. The central crossover membersextend between the first and second leg and between the second crossovermembers. The outermost crossover members extend for a firstcross-sectional area. The second crossover members extend for a secondcross-sectional area and the central crossover member extends for athird cross-sectional area. The first cross-sectional area is greaterthan the second cross-sectional area and the second cross-sectional areais greater than the third cross-sectional area.

In another embodiment according to the previous embodiment, a ratio ofthe first cross-sectional area to the second cross-sectional area to thethird cross-sectional area is 3:2:1.

In another embodiment according to any of the previous embodiments, atleast one of the leading and trailing edge is the leading edge.

In another embodiment according to any of the previous embodiments,there are intermediate connectors between the outermost crossover holesand the second crossover holes, and between the second crossover holesand the central crossover holes, and also between individual ones of thecentral crossover holes.

In another embodiment according to any of the previous embodiments, theconnectors extend for a cross-sectional area that is less than the firstand the second cross-sectional area.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 shows a first component.

FIG. 3 shows a molding intermediate step.

FIG. 4A shows a lost core.

FIG. 4B is a cross-sectional view along line B-B of FIG. 4A.

FIG. 5 shows another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (′TSFC)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

FIG. 2 shows a vane 100 which may be incorporated into the turbinesection of engine 20 of FIG. 1. Vane 100 includes an airfoil 106extending from a leading edge 108 to a trailing edge 110. The airfoil106 also extends between platforms 102 and 104.

A cooling channel 112 is formed adjacent the leading edge 108. While thecooling channel 112 is shown adjacent the leading edge 108, theteachings of this application may also extend to cooling channels formedadjacent the trailing edge 110.

A second cooling channel 113 is formed spaced from the leading edge 108relative to the channel 112. Crossover holes 116, 120, 124, 128 and 132communicate air from the channel 113 into the channel 112. Outlet holes111 may communicate the channel 112 through an outer skin of the airfoil106 for skin cooling.

Intermediate solid connectors 118, 122, 126 and 130 extend into an outerplane of FIG. 2 and provide a solid connection between opposed walls,not illustrated.

As can be appreciated from FIG. 2, the crossover hole 116 is of agreater dimension measured between the platforms 102 and 104 than theother crossover hole. This could be defined as a radial dimension alonga radial dimension R. The crossover hole 120 does not extend as far asthe hole 116, but extends for a greater cross-sectional area than theholes 124. Hole 120 extends for a cross-sectional area similar to thatof 128. Hole 116 may extend for a similar cross-sectional area as thatof hole 132.

In embodiments, the holes 116 and 132 may extend for three times thecross-sectional area of the holes 124. The holes 120 and 128 may extendfor twice the cross-sectional area of holes 124.

FIG. 3 shows an intermediate molded product 140. A lost core 143includes a first leg 144 and crossover members, or frames, 150connecting the first leg 144 to a second leg 148. As known, the lostcore will be leached away, leaving cooling channels. The dimension ofthe crossover members 150 is not to scale in this Figure, but, rather,can be best understood from FIGS. 4A and B (described below). Pins 146form the film cooling holes (such as 111) through the skin 142 of theairfoil. Additional lost core molds 160 form other cooling channels.Portions 152 intermediate the crossover members receive molten metalwhen the component is molded around the lost core elements.

Subsequent to the step of FIG. 3, the lost core 143 is leached awayleaving the hollow structure as shown in FIG. 2.

FIG. 4A shows the lost core 143. Leg 144 is connected to leg 148 by aplurality of crossover members. As known, the shape of the lost corewill form the shape of cooling channels in the eventual product. Theouter crossover members 164 and 182 could be called frames, in that,they extend for a greater cross-sectional area than do the more centralmembers 172, 174, etc. The second most inward members 178 and 168 extendfor a cross-sectional area intermediate that of member 164 and length172. Again, the three cross-sectional areas may be at the ratio of3:2:1, however, other relative cross-sectional areas would come withinthe scope of this application.

Hollows 162 in the core 143 receive metal at the outer ends of themolded part.

Hollows 180, 176, 170 and 166 receive molten metal to form theconnectors 118, 122, 126 and 130 as described above. Hollows extend fora cross-sectional area that is less than the first and secondcross-sectional area.

FIG. 4B is a cross-sectional view along line B-B of FIG. 4A. As shown,the members 164 and 182 extend for a greater cross-sectional area whichwill be a cross-sectional area than do the frames 168 and 178. Morecentral members 172 and 174 extend for even less of a cross-sectionalarea.

FIG. 5 shows a blade 200 having an airfoil 202 extending from a leadingedge 204 to a trailing edge 206. While FIGS. 2 and 3 show the teachingsof this application as formed in a static vane, it should be understoodthat the teachings would extend to a rotating blade 200 such as thatshown in FIG. 5.

The use of the several thicker frame members ensures that the mold core143 will be more rigid and less likely to break than the prior art. Asfurther known, the mold cores may be formed of an appropriate material.A worker of ordinary skill in the art would recognize the materialsgenerally utilized to form a lost core mold portion.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

The invention claimed is:
 1. A gas turbine engine component having anairfoil comprising: an airfoil extending between a leading edge and atrailing edge, and there being a first and a second cooling channel withsaid first cooling channel being spaced closest to one of said leadingand trailing edges and said second channel being spaced from said firstchannel relative to said one of said leading and trailing edges, andcrossover holes connecting said first and second cooling channels, withsaid plurality of crossover holes including outermost crossover holesspaced from each other at adjacent ends of each of said first and secondcooling channels, second crossover holes spaced closer to each otherthan are said outermost crossover holes, and central crossover holesbetween said second crossover holes; and said outermost crossover holesextending for a first cross-sectional area, said second crossover holesextending for a second cross-sectional area and said central crossoverholes extending for a third cross-sectional area, with said firstcross-sectional area being greater than said second cross-sectional areaand said second cross-sectional area being greater than said thirdcross-sectional area.
 2. The gas turbine engine component as set forthin claim 1, wherein a ratio of said first cross-sectional area to saidsecond cross-sectional area to said third cross-sectional area is 3:2:1.3. The gas turbine engine component as set forth in claim 2, whereinfilm cooling holes extend from said first cooling channel through a skinof said component.
 4. The gas turbine engine component as set forth inclaim 3, wherein said at least one of said leading and trailing edge issaid leading edge.
 5. The gas turbine engine component as set forth inclaim 4, wherein there are intermediate connectors between saidoutermost crossover holes and said crossover holes, and between saidsecond crossover holes and said central crossover holes, and alsobetween individual ones of said central crossover holes.
 6. The gasturbine engine component as set forth in claim 5, wherein saidintermediate connectors extend for a cross-sectional area that is lessthan said first and said second cross-sectional area.
 7. The gas turbineengine component as set forth in claim 1, wherein film cooling holesextend from said first cooling channel through a skin of said component.8. The gas turbine engine component as set forth in claim 1, whereinsaid at least one of said leading and trailing edge is said leadingedge.
 9. The gas turbine engine component as set forth in claim 1,wherein there are intermediate connectors between said outermostcrossover holes and said crossover holes, and between said secondcrossover holes and said central crossover holes, and also betweenindividual ones of said central crossover holes.
 10. The gas turbineengine component as set forth in claim 1, wherein said intermediateconnectors extend for a cross-sectional area that is less than saidfirst and said second cross-sectional area.
 11. A gas turbine enginecomprising: a compressor section and a turbine section, said turbinesection including rotors carrying rotating blades and static airfoilswith at least one of said rotating blades and said static airfoilsincluding an airfoil extending between a leading edge and a trailingedge, and there being a first and a second cooling channel with saidfirst cooling channel being spaced closest to one of said leading andtrailing edges and said second channel being spaced from said firstchannel relative to said one of said leading and trailing edges, andcrossover holes connecting said first and second cooling channels, withsaid plurality of crossover holes including outermost crossover holesspaced from each other at adjacent ends of each of said first and secondcooling channels, second crossover holes spaced closer to each otherthan are said outermost crossover holes, and central crossover holesbetween said second crossover holes; and said outermost crossover holesextending for a first cross-sectional area, said second crossover holesextending for a second cross-sectional area and said central crossoverholes extending for a third cross-sectional area, with said firstcross-sectional area being greater than said second cross-sectional areaand said second cross-sectional area being greater than said thirdcross-sectional area.
 12. The gas turbine engine as set forth in 11,wherein a ratio of said first cross-sectional area to said secondcross-sectional area to said third cross-sectional area is 3:2:1. 13.The gas turbine engine as set forth in claim 12, wherein said at leastone of said leading and trailing edge is said leading edge.
 14. The gasturbine engine as set forth in claim 13, wherein there are intermediateconnectors between said outermost crossover holes and said secondcrossover holes, and between said second crossover holes and saidcentral crossover holes, and also between individual ones of saidcentral crossover holes.
 15. The gas turbine engine as set forth inclaim 14, wherein said intermediate connectors extend for across-sectional area that is less than said first and said secondcross-sectional area.
 16. The gas turbine engine as set forth in claim11, wherein said at least one of said leading and trailing edge is saidleading edge.
 17. The gas turbine engine as set forth in claim 16,wherein there are intermediate connectors between said outermostcrossover holes and said second crossover holes, and between said secondcrossover holes and said central crossover holes, and also betweenindividual ones of said central crossover holes.
 18. The gas turbineengine as set forth in claim 17, wherein said intermediate connectorsextend for a cross-sectional area that is less than said first and saidsecond cross-sectional area.
 19. The gas turbine engine as set forth inclaim 11, wherein there are intermediate connectors between saidoutermost crossover holes and said second crossover holes, and betweensaid second crossover holes and said central crossover holes, and alsobetween individual ones of said central crossover holes.
 20. The gasturbine engine as set forth in claim 19, wherein said intermediateconnectors extend for a cross-sectional area that is less than saidfirst and said second cross-sectional area.